2.0 Analysis 2.1 Introduction The following analysis concentrates on the flight, the aircraft, the propeller, and the manufacturing of the blades. 2.2 Flight 2.2.1 Absence of Visual and Sensory Indications Because the point of origin of the fracture was inside the bore, the fatigue cracks propagated from the interior toward the blade surface. It was, therefore, impossible for the crew to detect any discrepancies during the external inspection of the aircraft. Additionally, there were no indications prior to or during the flight which could have enabled the flight crew to anticipate the fracture of the blade. Analysis of the FDR data for the occurrence flight and for previous flights revealed no failures or abnormal vibrations prior to the blade fracture. 2.2.2 Crew Performance Fracture of a propeller blade in flight is considered a highly improbable occurrence, as is the partial disintegration of an engine. Crews are not specifically trained to deal with such emergencies. In an emergency situation, the crew will react in accordance with the procedures practised during training. In the absence of training or standard procedures dictating how to deal with a particular situation, the crew must react by drawing upon their knowledge and experience. The crew reacted first to the cabin depressurization, then to the engine failure. The pilots were not fully aware of the situation until the flight attendant informed them that the forward portion of the engine had separated in flight and that there was a cut in the fuselage. When it became evident that they had indeed lost the right engine, the co-pilot cut off fuel to the engine by pulling fire handle No. 2. After stabilizing the aircraft and controlling the emergency, the crew assessed the situation. Because the behaviour of the aircraft in horizontal flight was satisfactory and the crew did not know the extent of the damage to the aircraft, the pilot-in-command felt it was preferable to minimize the number of turns in order to reduce the risk of further structural damage. The crew took into consideration the position of the aircraft, the weather, the airport services available, the known damage, the reactions of the aircraft, and the flying time to possible destinations. Analysis of ATS communications and the pilots' statements suggests that the decision to continue the flight to Montreal was taken nine minutes after the blade fracture. 2.3 Propeller Separation When the blade separated, the forces induced by the propeller imbalance on the three forward engine mounts exceeded the ultimate limits of each support and of the reduction gearbox mount. This allowed the propeller and reduction gearbox to separate from the turbine. 2.4 Fuselage Damage None of the propeller or reduction gearbox components found showed traces of paint from the fuselage. All indications are that the No. 2 blade fractured when in a position such that its trajectory allowed it to penetrate and then pass through the fuselage before following its course. The other damage to the aircraft was caused by nacelle debris during the separation. 2.5 Corrosion There were traces of chlorine in the corrosion pit at the point of origin of the No. 2 blade fracture. The chlorine was associated with the bleaching of the corks during their manufacture. Chlorine deposits on the surface of the cork, in the presence of water, produce an acidic solution which can attack the anodic coating on the aluminum and initiate superficial corrosion pitting. There were five corrosion pits in the taper bore of the No. 2 blade. Only one of the pits observed had progressed to the point where it initiated the fatigue fracture. This was the only pit on the face side of the blade bore. It is on this side of the blade bore, and at this station, that tensile stresses are greater than at any other point in the bore. 2.6 Blades The blades on this type of aircraft are manufactured in accordance with their Federal Aviation Administration (FAA) certification. The investigation revealed that shotpeening as part of the blade manufacturing process was discontinued in April 1987. Shotpeening offered supplementary protection in the form of reduced crack propagation in the area of the residual compressive stress induced by the shotpeening. Corks have been used for many years to hold lead wool in place in blade bores. Hamilton Standard had never observed that corrosion had been initiated by the use of corks. Additionally, the taper bore was not an area of the blade that was susceptible to corrosion. Internal visual inspection during major inspections was the only kind of inspection required by the manufacturer for this area. The moisture necessary to initiate the corrosion pitting must have been introduced when the cork was installed: that is, either at the 7,500-hour major inspection or during the manufacturing process. However, it could not be determined with certainty at which moment the moisture was introduced. 3.0 Conclusions 3.1 Findings The aircraft was certified, equipped and maintained in accordance with existing regulations and approved procedures. The blade was manufactured and inspected in accordance with the manufacturer's standards and procedures. The corks used in the taper bore are covered with a chlorine deposit. Water in the taper bore in contact with the chlorine from the cork can produce an acidic solution which can cause corrosion. The fractured blade showed corrosion pitting in the taper bore. The inside of the taper bore of the fractured blade had not been shotpeened during manufacture. An inspection of the No. 2 blade three days before the occurrence did not reveal any damage to the blade surface. The 7,500-hour major inspection was performed using the procedures and materials specified by the manufacturer. The fractured section of the blade punctured the fuselage and caused the cabin depressurization. 3.2 Causes Corrosion pitting had occurred on the surface of the taper bore of the No. 2 blade as a result of water combining with chlorine deposits on the cork in the taper bore. The chlorine was associated with the bleaching of the corks during their manufacture. One of the corrosion pits was the point of origin of the fatigue cracks that caused the propeller blade to fracture. 4.0 Safety Action 4.1 Action Taken The Board issued an Aviation Safety Advisory requesting that Transport Canada confirm with Hamilton Standard and the FAA that the measures taken by Hamilton Standard meet Canadian airworthiness requirements. Hamilton Standard was able to take the following steps to prevent similar occurrences: 1)Alert Service Bulletin 14SF-61-A73, dated 18 April 1994 (mandated by FAA Airworthiness Directive 94-09-06, effective date 02 May 1994), provided instructions to perform a one-time inspection of all 14SF blades using ultrasonic methods to inspect the taper bore for anomalies. 14SF-61-A74, dated 29 August 1994, provided instructions for the performance of a blade taper bore ultrasonic inspection as described above every 1,250 cycles, or to perform a one-time removal of the taper bore cork and visually inspect that portion of the taper bore for pits for blades that had not been shotpeened or had not had their taper bores inspected at Hamilton Standard's overhaul facility. The cork removal/visual inspection procedure is addressed by Service Bulletin 14SF-61-75. At the present time, the FAA is preparing an Airworthiness Directive to mandate this bulletin. 3)Revision No. 8 to the Hamilton Standard Component Maintenance Manual 61-13-02, dated 01 September 1994, includes instructions for inspection and repair of the blade taper bore area when the blade is returned to a repair facility. It also requires shotpeening of the taper bore for blades not previously shotpeened and mandates the use of new tools to reduce the chance of damaging the taper bore during lead wool removal. It also deletes instructions to install cork in the taper bore. 1)Re-introduced shotpeening to blade taper bores in May 1994. 2)Deleted cork installation in the taper bores in May 1994. 3)Changed the sequence of steps associated with the manufacturing to prevent water being introduced into the taper bore after lead wool is installed.